1. Field of the Invention
This invention relates to an improvement in the structure of a carbon-carbon nose tip for a re-entry space vehicle and more particularly, but not by way of limitation, to a sublimating thermal protection carbon-carbon nose tip structure which provides heat absorption externally and internally of the nose tip structure during re-entry and a preferred method of manufacture.
2. Description of the Prior Art
It is well known that an object entering the earth's atmosphere at even moderate speed is heated considerably by the attendant aerodynamic conditions at the surface of the object. Such heating occurs when vehicles re-enter the atmosphere and is a serious problem, forcing drastic limitations on the maximum re-entry speed which a structure of given design can attain without being severely damaged or even destroyed. This heat is primarily developed and is largely concentrated at the nose.
An important object of the present invention is to efficiently dissipate the heat developed at the nose. By achieving this object, the invention makes it possible to use higher re-entry speeds and therefore gives greater freedom in space vehicle design.
The nose tip of a re-entry vehicle is required to withstand heating encountered during atmospheric entry, maintain the structural integrity of the vehicle, prevent overheating of the payload, and maintain the aerodynamic characteristics of the vehicle. As the vehicle travels through the earth's atmosphere it experiences frictional heating in the boundary layer at its surface. The nose tip is also subjected to heat from gases that are at elevated temperatures as a result of being decelerated by the bow shock wave.
The amount of heat transferred to the nose tip depends on its shape and the materials from which it is made. The bow shock wave heats the gases behind it. The heat reaches the nose tip in the form of convection and radiation through the boundary layer adjacent to the surface.
Even for a properly designed nose tip shape, it is inevitable that some fraction of the vehicle's initial kinetic energy will finally reach the nose tip in the form of heat. Ablation is used to provide surface protection. Heat can be diverted from the re-entry space vehicle by allowing the nose tip's outer layer of material to melt, vaporize or sublime. While ablation provides excellent thermal protection, the resulting change in profile due to surface recession can adversely change the aerodynamic characteristics of the space vehicle. Additionally, adequate strength must be provided to prevent mechanical erosion of the nose tip by aerodynamic shear stresses.
Although the nose tip is a sacrificial item, it is desired that ablation be controlled, i.e., that the nose tip profile remain substantially the same throughout the period of re-entry, under any weather conditions, which may range from essentially clean air to high levels of dust and water droplets.
The manner in which an ablative material absorbs thermal energy is basically due to its ability to dissipate absorbed thermal energy by melting, vaporization and/or sublimation of surface material. By absorbing and dissipating large amounts of thermal heat the ablative material limits the temperature rise of the underlying structural shell and the internal components of the vehicle. While temperature limitation generally is the principal criterion in selecting and ablative material, mechanical performance of the ablative material generally is also an important consideration.
Lindberg, Jr., U.S. Pat. No. 3,682,100 discloses in a space vehicle a nose having an imperforate outer surface and incorporating adjacent such surface a heat-dissociable metallic hydride mixed within a porous ceramic so that heat developed at the nose surface during flight in the atmosphere is used to cause the hydride to emit hydrogen gas, thereby tending to reduce the temperature of the nose surface.
Taverna et al, U.S. Pat. No. 4,515,847 discloses a re-entry carbon-carbon nose tip structure which has its outer portion loaded with a heat-resisting particulate material to provide a desired ablative performance in said nose tip structure.
Moores et al, U.S. Pat. No. 4,131,708 discloses an ablative carbon-composite shaped structure for high temperature thermal protection which includes a carbon-carbon composite reinforced body which is provided with implants in the form of elongate columns of a refractory metal carbide that are aligned parallel with one another axially in the direction of expected thermal flux and which are positioned at predetermined locations of expected thermal flux through said structure.
It is apparent that the prior art does not disclose a carbon-carbon space structure which uses a first sublimation activity on its outer surface when very high heat fluxes are encountered during re-entry and a second sublimation activity on its inner surface when the internal surface reaches a specific temperature thereby enabling the space structure to retain its basic geometry while sustaining a thermal environment well in excess of that presently achievable by known thermal protection systems. While certain aspects of the above disclosed prior art are of interest, they do not teach the particular thermal protection system utilizing two independent sublimating activities nor do they teach the subject method of fabricating such a novel thermal protection system for a carbon-carbon structure for a re-entry vehicle.